Flared central cavity aft of airfoil leading edge

ABSTRACT

A blade includes an airfoil defined by a pressure side outer wall and a suction side outer wall connecting along leading and trailing edges and form a radially extending chamber for receiving a coolant flow. A rib configuration may include: a leading edge transverse rib connecting the pressure side outer wall and the suction side outer wall and partitioning the radially extending chamber into a leading edge passage within the leading edge of the airfoil and a central passage adjacent to the leading edge passage. One or both camber line ribs connect to a corresponding pressure side outer wall and suction side outer wall at a point aft of the leading edge transverse rib causing the central passage to extend towards one or both of the pressure side outer wall and the suction side outer wall, resulting in a flared center cavity aft of the leading edge.

BACKGROUND OF THE INVENTION

This disclosure relates to turbine airfoils, and more particularly tohollow turbine airfoils, such as rotor or stator blades, having internalchannels for passing fluids such as air to cool the airfoils.

Combustion or gas turbine engines (hereinafter “gas turbines”) include acompressor, a combustor, and a turbine. As is well known in the art, aircompressed in the compressor is mixed with fuel and ignited in thecombustor and then expanded through the turbine to produce power. Thecomponents within the turbine, particularly the circumferentiallyarrayed rotor and stator blades, are subjected to a hostile environmentcharacterized by the extremely high temperatures and pressures of thecombustion products that are expended therethrough. In order towithstand the repetitive thermal cycling as well as the extremetemperatures and mechanical stresses of this environment, the airfoilsmust have a robust structure and be actively cooled.

As will be appreciated, turbine rotor and stator blades often containinternal passageways or circuits that form a cooling system throughwhich a coolant, typically air bled from the compressor, is circulated.Such cooling circuits are typically formed by internal ribs that providethe required structural support for the airfoil, and include multipleflow path arrangements to maintain the airfoil within an acceptabletemperature profile. The air passing through these cooling circuitsoften is vented through film cooling apertures formed on the leadingedge, trailing edge, suction side, and pressure side of the airfoil.

It will be appreciated that the efficiency of gas turbines increases asfiring temperatures rise. Because of this, there is a constant demandfor technological advances that enable turbine blades to withstand everhigher temperatures. These advances sometimes include new materials thatare capable of withstanding the higher temperatures, but just as oftenthey involve improving the internal configuration of the airfoil so toenhance the blades structure and cooling capabilities. However, becausethe use of coolant decreases the efficiency of the engine, newarrangements that rely too heavily on increased levels of coolant usagemerely trade one inefficiency for another. As a result, there continuesto be demand for new airfoil arrangements that offer internal airfoilconfigurations and coolant circulation that improves coolant efficiency.

A consideration that further complicates arrangement of internallycooled airfoils is the temperature differential that develops duringoperation between the airfoils internal and external structure. That is,because they are exposed to the hot gas path, the external walls of theairfoil typically reside at much higher temperatures during operationthan many of the internal ribs, which, for example, may have coolantflowing through passageways defined to each side of them. In fact, acommon airfoil configuration includes a “four-wall” arrangement in whichlengthy inner ribs run parallel to the pressure and suction side outerwalls. It is known that high cooling efficiency can be achieved by thenear-wall flow passages that are formed in the four-wall arrangement. Achallenge with the near-wall flow passages is that the outer wallsexperience a significantly greater level of thermal expansion than theinner walls. This imbalanced growth causes stress to develop at thepoints at which the inner ribs connect, which may cause low cyclicfatigue that can shorten the life of the blade.

BRIEF DESCRIPTION OF THE INVENTION

A first aspect of the disclosure provides a blade comprising an airfoildefined by a concave pressure side outer wall and a convex suction sideouter wall that connect along leading and trailing edges and,therebetween, form a radially extending chamber for receiving the flowof a coolant, the blade further comprising: a rib configurationincluding: a leading edge transverse rib connecting the pressure sideouter wall and the suction side outer wall and partitioning the radiallyextending chamber into a leading edge passage within the leading edge ofthe airfoil and a central passage adjacent to the leading edge passage,and a camber line rib connecting to a selected one of the pressure sideouter wall and the suction side outer wall at a point aft of the leadingedge transverse rib causing the central passage to extend towards theselected one of the pressure side outer wall and the suction side outerwall.

A second aspect of the disclosure provides a turbine rotor bladecomprising an airfoil defined by a concave pressure side outer wall anda convex suction side outer wall that connect along leading and trailingedges and, therebetween, form a radially extending chamber for receivingthe flow of a coolant, the turbine rotor blade further comprising: a ribconfiguration including: a leading edge transverse rib connecting thepressure side outer wall and the suction side outer wall andpartitioning the radially extending chamber into a leading edge passagewithin the leading edge of the airfoil and a central passage adjacent tothe leading edge passage, and a camber line rib connecting to a selectedone of the pressure side outer wall and the suction side outer wall at apoint aft of the leading edge transverse rib causing the central passageto extend towards the selected one of the pressure side outer wall andthe suction side outer wall.

The illustrative aspects of the present disclosure are arrangements tosolve the problems herein described and/or other problems not discussed.

BRIEF DESCRIPTION OF THE DRAWINGS

These and other features of this disclosure will be more readilyunderstood from the following detailed description of the variousaspects of the disclosure taken in conjunction with the accompanyingdrawings that depict various embodiments of the disclosure, in which:

FIG. 1 is a schematic representation of an example turbine engine inwhich certain embodiments of the present application may be used.

FIG. 2 is a sectional view of the compressor section of the combustionturbine engine of FIG. 1.

FIG. 3 is a sectional view of the turbine section of the combustionturbine engine of FIG. 1.

FIG. 4 is a perspective view of a turbine rotor blade of the type inwhich embodiments of the present disclosure may be employed.

FIG. 5 is a cross-sectional view of a turbine rotor blade having aninner wall or rib configuration according to conventional arrangement.

FIG. 6 is a cross-sectional view of a turbine rotor blade having aninner wall configuration according to conventional arrangement.

FIG. 7 is a cross-sectional view of a turbine rotor blade having aflared central passage configuration according to an alternativeembodiment of the present disclosure.

FIG. 8 is a cross-sectional view of a turbine rotor blade having aflared central passage configuration according to an alternativeembodiment of the present disclosure.

FIG. 9 is a cross-sectional view of a turbine rotor blade having aflared central passage configuration according to an alternativeembodiment of the present disclosure.

FIG. 10 is a cross-sectional view of a turbine rotor blade having aflared central passage without a wavy profile as in FIGS. 7-9, accordingto an alternative embodiment of the present disclosure.

It is noted that the drawings of the disclosure are not necessarily toscale. The drawings are intended to depict only typical aspects of thedisclosure, and therefore should not be considered as limiting the scopeof the disclosure. In the drawings, like numbering represents likeelements between the drawings.

DETAILED DESCRIPTION OF THE INVENTION

As an initial matter, in order to clearly describe the currentdisclosure it will become necessary to select certain terminology whenreferring to and describing relevant machine components within a gasturbine. When doing this, if possible, common industry terminology willbe used and employed in a manner consistent with its accepted meaning.Unless otherwise stated, such terminology should be given a broadinterpretation consistent with the context of the present applicationand the scope of the appended claims. Those of ordinary skill in the artwill appreciate that often a particular component may be referred tousing several different or overlapping terms. What may be describedherein as being a single part may include and be referenced in anothercontext as consisting of multiple components. Alternatively, what may bedescribed herein as including multiple components may be referred toelsewhere as a single part.

In addition, several descriptive terms may be used regularly herein, andit should prove helpful to define these terms at the onset of thissection. These terms and their definitions, unless stated otherwise, areas follows. As used herein, “downstream” and “upstream” are terms thatindicate a direction relative to the flow of a fluid, such as theworking fluid through the turbine engine or, for example, the flow ofair through the combustor or coolant through one of the turbine'scomponent systems. The term “downstream” corresponds to the direction offlow of the fluid, and the term “upstream” refers to the directionopposite to the flow. The terms “forward” and “aft”, without any furtherspecificity, refer to directions, with “forward” referring to the frontor compressor end of the engine, and “aft” referring to the rearward orturbine end of the engine. It is often required to describe parts thatare at differing radial positions with regard to a center axis. The term“radial” refers to movement or position perpendicular to an axis. Incases such as this, if a first component resides closer to the axis thana second component, it will be stated herein that the first component is“radially inward” or “inboard” of the second component. If, on the otherhand, the first component resides further from the axis than the secondcomponent, it may be stated herein that the first component is “radiallyoutward” or “outboard” of the second component. The term “axial” refersto movement or position parallel to an axis. Finally, the term“circumferential” refers to movement or position around an axis. It willbe appreciated that such terms may be applied in relation to the centeraxis of the turbine.

By way of background, referring now to the figures, FIGS. 1 through 4illustrate an example combustion turbine engine in which embodiments ofthe present application may be used. It will be understood by thoseskilled in the art that the present disclosure is not limited to thisparticular type of usage. The present disclosure may be used incombustion turbine engines, such as those used in power generation,airplanes, as well as other engine or turbomachine types. The examplesprovided are not meant to be limiting unless otherwise stated.

FIG. 1 is a schematic representation of a combustion turbine engine 10.In general, combustion turbine engines operate by extracting energy froma pressurized flow of hot gas produced by the combustion of a fuel in astream of compressed air. As illustrated in FIG. 1, combustion turbineengine 10 may be configured with an axial compressor 11 that ismechanically coupled by a common shaft or rotor to a downstream turbinesection or turbine 13, and a combustor 12 positioned between compressor11 and turbine 13.

FIG. 2 illustrates a view of an illustrative multi-staged axialcompressor 11 that may be used in the combustion turbine engine ofFIG. 1. As shown, compressor 11 may include a plurality of stages. Eachstage may include a row of compressor rotor blades 14 followed by a rowof compressor stator blades 15. Thus, a first stage may include a row ofcompressor rotor blades 14, which rotate about a central shaft, followedby a row of compressor stator blades 15, which remain stationary duringoperation.

FIG. 3 illustrates a partial view of an illustrative turbine section orturbine 13 that may be used in the combustion turbine engine of FIG. 1.Turbine 13 may include a plurality of stages. Three illustrative stagesare shown, but more or less stages may be present in the turbine 13. Afirst stage includes a plurality of turbine buckets or turbine rotorblades 16, which rotate about the shaft during operation, and aplurality of nozzles or turbine stator blades 17, which remainstationary during operation. Turbine stator blades 17 generally arecircumferentially spaced one from the other and fixed about the axis ofrotation. Turbine rotor blades 16 may be mounted on a turbine wheel (notshown) for rotation about the shaft (not shown). A second stage ofturbine 13 also is illustrated. The second stage similarly includes aplurality of circumferentially spaced turbine stator blades 17 followedby a plurality of circumferentially spaced turbine rotor blades 16,which are also mounted on a turbine wheel for rotation. A third stagealso is illustrated, and similarly includes a plurality of turbinestator blades 17 and rotor blades 16. It will be appreciated thatturbine stator blades 17 and turbine rotor blades 16 lie in the hot gaspath of the turbine 13. The direction of flow of the hot gases throughthe hot gas path is indicated by the arrow. As one of ordinary skill inthe art will appreciate, turbine 13 may have more, or in some casesless, stages than those that are illustrated in FIG. 3. Each additionalstage may include a row of turbine stator blades 17 followed by a row ofturbine rotor blades 16.

In one example of operation, the rotation of compressor rotor blades 14within axial compressor 11 may compress a flow of air. In combustor 12,energy may be released when the compressed air is mixed with a fuel andignited. The resulting flow of hot gases from combustor 12, which may bereferred to as the working fluid, is then directed over turbine rotorblades 16, the flow of working fluid inducing the rotation of turbinerotor blades 16 about the shaft. Thereby, the energy of the flow ofworking fluid is transformed into the mechanical energy of the rotatingblades and, because of the connection between the rotor blades and theshaft, the rotating shaft rotates. The mechanical energy of the shaftmay then be used to drive the rotation of the compressor rotor blades14, such that the necessary supply of compressed air is produced, andalso, for example, a generator to produce electricity.

FIG. 4 is a perspective view of a turbine rotor blade 16 of the type inwhich embodiments of the present disclosure may be employed. Turbinerotor blade 16 includes a root 21 by which rotor blade 16 attaches to arotor disc. Root 21 may include a dovetail configured for mounting in acorresponding dovetail slot in the perimeter of the rotor disc. Root 21may further include a shank that extends between the dovetail and aplatform 24, which is disposed at the junction of airfoil 25 and root 21and defines a portion of the inboard boundary of the flow path throughturbine 13. It will be appreciated that airfoil 25 is the activecomponent of rotor blade 16 that intercepts the flow of working fluidand induces the rotor disc to rotate. While the blade of this example isa turbine rotor blade 16, it will be appreciated that the presentdisclosure also may be applied to other types of blades within turbineengine 10, including turbine stator blades 17 (vanes). It will be seenthat airfoil 25 of rotor blade 16 includes a concave pressure side (PS)outer wall 26 and a circumferentially or laterally opposite convexsuction side (SS) outer wall 27 extending axially between oppositeleading and trailing edges 28, 29 respectively. Sidewalls 26 and 27 alsoextend in the radial direction from platform 24 to an outboard tip 31.(It will be appreciated that the application of the present disclosuremay not be limited to turbine rotor blades, but may also be applicableto stator blades. The usage of rotor blades in the several embodimentsdescribed herein is illustrative unless otherwise stated.)

FIGS. 5 and 6 show two example internal wall constructions as may befound in a rotor blade airfoil 25 having a conventional arrangement. Asindicated, an outer surface of airfoil 25 may be defined by a relativelythin pressure side (PS) outer wall 26 and suction side (SS) outer wall27, which may be connected via a plurality of radially extending andintersecting ribs 60. Ribs 60 are configured to provide structuralsupport to airfoil 25, while also defining a plurality of radiallyextending and substantially separated flow passages 40. Typically, ribs60 extend radially so to partition flow passages 40 over much of theradial height of airfoil 25, but the flow passage may be connected alongthe periphery of the airfoil so to define a cooling circuit. That is,flow passages 40 may fluidly communicate at the outboard or inboardedges of airfoil 25, as well as via impingement apertures (latter notshown) that may be positioned therebetween. In this manner certain offlow passages 40 together may form a winding or serpentine coolingcircuit. Additionally, film cooling ports (not shown) may be includedthat provide outlets through which coolant is released from flowpassages 40 onto outer surface of airfoil 25.

Ribs 60 may include two different types, which then, as provided herein,may be subdivided further. A first type, a camber line rib 62, istypically a lengthy rib that extends in parallel or approximatelyparallel to the camber line of the airfoil, which is a reference linestretching from a leading edge 28 to a trailing edge 29 that connectsthe midpoints between pressure side outer wall 26 and suction side outerwall 27. As is often the case, the illustrative conventionalconfiguration of FIGS. 5 and 6 include two camber line ribs 62, apressure side camber line rib 63, which also may be referred to as thepressure side outer wall given the manner in which it is offset from andclose to the pressure side outer wall 26, and a suction side camber linerib 64, which also may be referred to as the suction side outer wallgiven the manner in which it is offset from and close to the suctionside outer wall 27. As mentioned, these types of arrangements are oftenreferred to as having a “four-wall” configuration due to the prevalentfour main walls that include two outer walls 26, 27 and two camber lineribs 63, 64. It will be appreciated that outer walls 26, 27 and camberline ribs 62 may be formed using any now known or later developedtechnique, e.g., via casting or additive manufacturing as integralcomponents.

The second type of rib is referred to herein as a transverse rib 66.Transverse ribs 66 are the shorter ribs that are shown connecting thewalls and inner ribs of the four-wall configuration. As indicated, thefour walls may be connected by a number of transverse ribs 66, which maybe further classified according to which of the walls each connects. Asused herein, traverse ribs 66 that connect pressure side outer wall 26to pressure side camber line rib 63 are referred to as pressure sidetraverse ribs 67. Transverse ribs 66 that connect suction side outerwall 27 to suction side camber line rib 64 are referred to as suctionside traverse ribs 68. Transverse ribs 66 that connect pressure sidecamber line rib 63 to suction side camber line rib 64 are referred to ascenter traverse ribs 69. Finally, a transverse rib 66 that connectspressure side outer wall 26 and suction side outer wall 27 near leadingedge 28 is referred to as a leading edge transverse rib 70. Leading edgetransverse rib 70, in FIGS. 5 and 6, also connects to a leading edge endof pressure side camber line rib 63 and a leading edge end of suctionside camber line rib 64.

As leading edge transverse rib 70 couples pressure side outer wall 26and suction side outer wall 27, it also forms passage 40 referred toherein as a leading edge passage 42. Leading edge passage 42 may havesimilar functionality as other passages 40, described herein.

In general, the purpose of any internal configuration in an airfoil 25is to provide efficient near-wall cooling, in which the cooling airflows in channels adjacent to outer walls 26, 27 of airfoil 25. It willbe appreciated that near-wall cooling is advantageous because thecooling air is in close proximity of the hot outer surfaces of theairfoil, and the resulting heat transfer coefficients are high due tothe high flow velocity achieved by restricting the flow through narrowchannels. However, such arrangements are prone to experiencing low cyclefatigue due to differing levels of thermal expansion experienced withinairfoil 25, which, ultimately, may shorten the life of the rotor blade.For example, in operation, suction side outer wall 27 thermally expandsmore than suction side camber line rib 64. This differential expansiontends to increase the length of the camber line of airfoil 25, and,thereby, causes stress between each of these structures as well as thosestructures that connect them. In addition, pressure side outer wall 26also thermally expands more than the cooler pressure side camber linerib 63. In this case, the differential tends to decrease the length ofthe camber line of airfoil 25, and, thereby, cause stress between eachof these structures as well as those structures that connect them. Theoppositional forces within the airfoil that, in the one case, tends todecrease the airfoil camber line and, in the other, increase it, canlead to stress concentrations. The various ways in which these forcesmanifest themselves given an airfoil's particular structuralconfiguration and the manner in which the forces are then balanced andcompensated for becomes a significant determiner of the part life ofrotor blade 16.

More specifically, in a common scenario, suction side outer wall 27tends to bow outward at the apex of its curvature as exposure to thehigh temperatures of the hot gas path cause it to thermally expand. Itwill be appreciated that suction side camber line rib 64, being aninternal wall, does not experience the same level of thermal expansionand, therefore, does not have the same tendency to bow outward. That is,camber line rib 64 and transverse ribs 66 and their connection pointsresists the thermal growth of the outer wall 27.

Conventional arrangements, an example of which is shown in FIG. 5, havecamber line ribs 62 formed with stiff geometries that provide little orno compliance. The resistance and the stress concentrations that resultfrom it can be substantial. Exacerbating the problem, traverse ribs 66used to connect camber line rib 62 to outer wall 27 may be formed withlinear profiles and generally oriented at right angles in relation tothe walls that they connect. This being the case, traverse ribs 66operated to basically hold fast the “cold” spatial relationship betweenthe outer wall 27 and the camber line rib 64 as the heated structuresexpand at significantly different rates. The little or no “give”situation prevents defusing the stress that concentrates in certainregions of the structure. The differential thermal expansion results inlow cycle fatigue issues that shorten component life.

Many different internal airfoil cooling systems and structuralconfigurations have been evaluated in the past, and attempts have beenmade to rectify this issue. One such approach proposes overcooling outerwalls 26, 27 so that the temperature differential and, thereby, thethermal growth differential are reduced. It will be appreciated, though,that the way in which this is typically accomplished is to increase theamount of coolant circulated through the airfoil. Because coolant istypically air bled from the compressor, its increased usage has anegative impact on the efficiency of the engine and, thus, is a solutionthat is preferably avoided. Other solutions have proposed the use ofimproved fabrication methods and/or more intricate internal coolingconfigurations that use the same amount of coolant, but use it moreefficiently. While these solutions have proven somewhat effective, eachbrings additional cost to either the operation of the engine or themanufacture of the part, and does nothing to directly address the rootproblem, which is the geometrical deficiencies of conventionalarrangement in light of how airfoils grow thermally during operation. Asshown in one example in FIG. 6, another approach employs certain curvingor bubbled or sinusoidal or wavy internal ribs (hereinafter “wavy ribs”)that alleviate imbalanced thermal stresses that often occur in theairfoil of turbine blades. These structures reduce the stiffness of theinternal structure of airfoil 25 so to provide targeted flexibility bywhich stress concentrations are dispersed and strain off-loaded to otherstructural regions that are better able to withstand it. This mayinclude, for example, off-loading stress to a region that spreads thestrain over a larger area, or, perhaps, structure that offloads tensilestress for a compressive load, which is typically more preferable. Inthis manner, life-shortening stress concentrations and strain may beavoided.

However, despite the above arrangements, a high stress area may stillresult at leading edge transverse rib 70 connection points 80 to camberline ribs 63 and 64, e.g., because camber line ribs 63, 64 load pathreacts at connection points 80 where insufficient cooling occurs

FIGS. 7-10 provide cross-sectional views of a turbine rotor blade 16having an inner wall or rib configuration according to embodiments ofthe present disclosure. Configuration of ribs that are typically used asboth structural support as well as partitions that divide hollowairfoils 25 into substantially separated radially extending flowpassages 40 that may be interconnects as desired to create coolingcircuits. These flow passages 40 and the circuits they form are used todirect a flow of coolant through the airfoil 25 in a particular mannerso that its usage is targeted and more efficient. Though the examplesprovided herein are shown as they might be used in a turbine rotorblades 16, it will be appreciated that the same concepts also may beemployed in turbine stator blades 17 (vanes).

Specifically, as will be described relative to FIGS. 7-10, the presentdisclosure teaches positioning at least one camber line rib 63, 64 toconnect to pressure side outer wall 26 and/or suction side outer wall 27at a point aft of leading edge transverse rib 70 causing central passage46 to extend towards pressure side outer wall 26 and/or suction sideouter wall 27. In this fashion, central passage 46 flares outwardlytoward at least one of outer walls 26, 27, relieving stress inconnection points 80 and other adjacent structure to leading edgetransverse rib 70. To affect this change, turbine rotor blade 16includes a rib configuration including a leading edge transverse rib 70connecting pressure side outer wall 26 and suction side outer wall 27and partitioning the radially extending chamber into a leading edgepassage 42 within leading edge 28 of airfoil 25 and a central passage 46adjacent to leading edge passage 42. Central passage 46 is ‘central’because it is within or surrounded by other passages, e.g., 48, 50,formed between camber line ribs 63, 64 and outer walls 26, 27,respectively. For example, pressure side outer wall 26 and pressure sidecamber line rib 63 define a pressure side flow passage 48 therebetweenand suction side outer wall 27 and suction side camber line rib 64define a suction side flow passage 50 therebetween.

As illustrated, as an option, crossover passage(s) 44 may be providedwithin leading edge transverse rib 70 to allow coolant to flow betweenleading edge passage 42 and immediately aft and adjacent central passage46. More specifically, as illustrated in FIG. 7, according toembodiments, a crossover passage 44 may allow coolant to pass to and/orfrom leading edge passage 42 to immediately aft central passage 46.Cross-over port 44 may include any number thereof positioned in aradially spaced relation between passages 40, 42. The stress describedherein that is created by camber line ribs 62, 63, 64 at points 80(FIGS. 5 and 6) may be more intense where crossover passages 44 areemployed between leading edge passage 42 and immediately aft centralpassage 46. In particular, where cross-over passages 44 are provided,camber line ribs 62, 63, 64 load path may react on connection points 80(FIGS. 5 and 6) where crossover passages 44 would be located, causinghigher stress. Crossover passage(s) 44 are not necessary in allembodiments, e.g., although applicable to the embodiment shown therein,FIG. 9 shows an example without crossover passage(s) 44. Where crossoverpassage(s) 44 are provided, however, the teachings of the disclosurerelieve stress adjacent thereto in leading edge transverse rib 70 andadjacent structure.

The rib configuration also includes a camber line rib 63, 64 connectingto a selected one of the pressure side outer wall 26 and the suctionside outer wall 27 at a point 92 aft of leading edge transverse rib 70causing central passage 46 to extend towards the selected one ofpressure side outer wall 26 and suction side outer wall 27. A camberline rib 62, as described above, is one of the longer ribs thattypically extend from a position near leading edge 28 of airfoil 25toward trailing edge 29. These ribs are referred to as “camber lineribs” because the path they trace is approximately parallel to thecamber line of airfoil 25, which is a reference line extending betweenleading edge 28 and trailing edge 29 of airfoil 25 through a collectionof points that are equidistant between concave pressure side outer wall26 and convex suction side outer wall 27. As shown, the ribconfiguration according to embodiments of the disclosure forms a flaredportion 90 that flares towards outer wall(s) 26, 27 in central cavity46. Since more coolant is flowing near leading edge transverse rib 70and crossover passage(s) 44 (where provided), the stress therein isreduced.

In one embodiment, shown in FIGS. 7-9, the rib configuration of thepresent disclosure includes camber line rib 62 having a wavy profile, asdescribed in US Patent Publication 2015/0184519, which is herebyincorporated by reference. (As used herein, the term “profile” isintended to refer to the shape the ribs have in the cross-sectionalviews of FIGS. 7-10.) According to the present application, a “wavyprofile” includes one that is noticeably curved and sinusoidal in shape,as indicated. In other words, the “wavy profile” is one that presents aback-and-forth “S” profile. In another embodiment, as shown in FIG. 10,the rib configuration of the present disclosure may include camber lineribs 63, 64 having a non-wavy profile.

With further reference to FIG. 7, according to one embodiment, bothpressure side camber line 63 and suction side camber line 64 connect toa respective outer wall 26, 27, i.e., the selected one outer wallincludes both (two) outer walls. That is, pressure side camber line rib63 residing near pressure side outer wall 26 connects to pressure sideouter wall 26 at a point 92 aft of leading edge transverse rib 70. Thisrib configuration causes central passage 46 to extend towards pressureside outer wall 26 with a first flared region 90. Further, suction sidecamber line rib 64 residing near suction side outer wall 27 connects tosuction side outer wall 27 at a point 92 aft of leading edge transverserib 70. This rib arrangement causes central passage 46 to extend towardssuction side outer wall 27 with a second flared region 90. Asillustrated, flared regions 90 include rounded interiors. In oneembodiment, at a location where central passage 46 extends towards outerwalls 26, 27, i.e., where flared regions 90 are present, a width t_(f)is defined between outermost extents of flared regions 90 (FIG. 7). Atthe same location (i.e., cross-sectional line as width t_(f) ismeasured), a width between outer surfaces of outer walls 26, 27 may bedefined as t_(a). A ratio of width t_(f) to width t_(a) (t_(f)/t_(a))may range from 40-70%, depending on desired configuration and embodimentof flared regions 90 employed. In an alternative embodiment, flaredregions 90 may have an outermost extent that is in line with anoutermost extent of adjacent passages 48, 50 relative to outer walls 26,27, i.e., outer walls 26, 27 have a thickness relative to respectiveflared regions 90 that is equal to or nearly equal to the thicknessthereof relative to corresponding adjacent passages 48, 50. In either ofthese embodiments, both leading edge transverse rib 70 and pressure sidecamber line rib 63 connect to pressure side outer wall 26 in a spacedmanner, and both leading edge transverse rib 70 and suction side camberline rib 64 connect to suction side outer wall 27 in a spaced manner.

In contrast, in alternative embodiments shown in FIGS. 8 and 9, only aselected one of the camber line ribs 63, 64 connects to a respectiveouter wall 26, 27. In FIG. 8, camber line rib 62 includes pressure sidecamber line rib 63 connected to pressure side outer wall 26 at point 92aft of leading edge transverse rib 70 causing central passage 46 toextend towards pressure side outer wall 26. Here, leading edgetransverse rib 70 and pressure side camber line rib 63 connect topressure side outer wall 26 in a spaced manner, but leading edgetransverse rib 70 and suction side camber line rib 64 connect to oneanother. Only one flared region 90 is present in this embodiment. Incontrast, in FIG. 9, camber line rib 62 includes suction side camberline rib 64 connected to suction side outer wall 27 at a point 92 aft ofleading edge transverse rib 70 causing central passage 46 to extendtowards suction side outer wall 27. In this embodiment, leading edgetransverse rib 70 and suction side camber line rib 64 connect to suctionside outer wall 27 in a spaced manner, but leading edge transverse rib70 and pressure side camber line rib 63 connect to one another. Again,only one flared region 90 is present here. Here again, flared regions 90include rounded interiors. Further, in one embodiment, flared regions 90may have a width t_(f) from an outermost extent of one flared region 90to an inner extent of central chamber 46. At the same cross-sectionalline as width t_(f) is measured, a width of outer walls 26, 27 (outersurfaces thereof) may be denoted t_(a). A ratio of width t_(f) to widtht_(a) (t_(f)/t_(a)) may range from 40-70%, depending on desiredconfiguration and embodiment of flared regions 90 employed. In analternative embodiment, flared regions 90 may have an outermost extentthat is in line with an outermost extent of adjacent passages 48, 50relative to outer walls 26, 27, i.e., outer walls 26, 27 have athickness relative to respective flared regions 90 that is equal to ornearly equal to the thickness thereof relative to corresponding adjacentpassages 48, 50. On a side that does not include flared regions 90, theadjacent passage 48 or 50 can be extended more toward transverse rib 70,compared to those shown in FIGS. 5 and 6. That is, adjacent passage 48or 50 is extended forwardly such that transverse rib 70 and an end of arespective camber line rib 63 or 64 coupling to outer wall 26 or 27collectively have a thickness equal to or nearly equal to that of otherribs 60, 62, 63, etc.—in contrast to thicker transverse ribs 63, 70 or64, 70 shown FIGS. 5 and 6.

FIG. 10 shows an alternative embodiment, similar to FIG. 7, except thatit does not employ a wavy profile. It is emphasized that the teachingsof FIGS. 8 and 9 may also be employed to rib configurations having anon-wavy profile. Further, the teachings of the disclosure may beapplied to a wide variety of rib configurations having leading edgepassage 42 and immediately aft central passage 46, as described herein.

The terminology used herein is for the purpose of describing particularembodiments only and is not intended to be limiting of the disclosure.As used herein, the singular forms “a”, “an” and “the” are intended toinclude the plural forms as well, unless the context clearly indicatesotherwise. It will be further understood that the terms “comprises”and/or “comprising,” when used in this specification, specify thepresence of stated features, integers, steps, operations, elements,and/or components, but do not preclude the presence or addition of oneor more other features, integers, steps, operations, elements,components, and/or groups thereof “Optional” or “optionally” means thatthe subsequently described event or circumstance may or may not occur,and that the description includes instances where the event occurs andinstances where it does not.

Approximating language, as used herein throughout the specification andclaims, may be applied to modify any quantitative representation thatcould permissibly vary without resulting in a change in the basicfunction to which it is related. Accordingly, a value modified by a termor terms, such as “about”, “approximately” and “substantially”, are notto be limited to the precise value specified. In at least someinstances, the approximating language may correspond to the precision ofan instrument for measuring the value. Here and throughout thespecification and claims, range limitations may be combined and/orinterchanged, such ranges are identified and include all the sub-rangescontained therein unless context or language indicates otherwise.“Approximately” as applied to a particular value of a range applies toboth values, and unless otherwise dependent on the precision of theinstrument measuring the value, may indicate +/−10% of the statedvalue(s).

The corresponding structures, materials, acts, and equivalents of allmeans or step plus function elements in the claims below are intended toinclude any structure, material, or act for performing the function incombination with other claimed elements as specifically claimed. Thedescription of the present disclosure has been presented for purposes ofillustration and description, but is not intended to be exhaustive orlimited to the disclosure in the form disclosed. Many modifications andvariations will be apparent to those of ordinary skill in the artwithout departing from the scope and spirit of the disclosure. Theembodiment was chosen and described in order to best explain theprinciples of the disclosure and the practical application, and toenable others of ordinary skill in the art to understand the disclosurefor various embodiments with various modifications as are suited to theparticular use contemplated.

What is claimed is:
 1. A blade comprising an airfoil defined by aconcave pressure side outer wall and a convex suction side outer wallthat connect along leading and trailing edges and, therebetween, form aradially extending chamber for receiving a flow of a coolant, the bladefurther comprising: a rib configuration including: a leading edgetransverse rib connecting the pressure side outer wall and the suctionside outer wall and partitioning the radially extending chamber into aleading edge passage within the leading edge of the airfoil and acentral passage adjacent to the leading edge passage; and a camber linerib connecting to a selected one of the pressure side outer wall or thesuction side outer wall at a point aft of the leading edge transverserib causing the central passage to extend towards the selected one ofthe pressure side outer wall or the suction side outer wall.
 2. Theblade of claim 1, wherein the camber line rib includes: a pressure sidecamber line rib residing near the pressure side outer wall and connectedto the pressure side outer wall at a point aft of the leading edgetransverse rib causing the central passage to extend towards thepressure side outer wall; and a suction side camber line rib residingnear the suction side outer wall and connected to the suction side outerwall at a point aft of the leading edge transverse rib causing thecentral passage to extend towards the suction side outer wall.
 3. Theblade of claim 2, wherein, at a location where the central passageextends towards the suction side outer wall and the pressure side outerwall, a width t_(f) is defined between outermost extents of the centralpassage and a width t_(a) is defined between outer surfaces of thesuction side outer wall and the pressure side outer wall, and wherein aratio of the width t_(f) to the width t_(a) ranges from 40% to 70%. 4.The blade of claim 2, wherein the pressure side outer wall and thepressure side camber line rib define a pressure side flow passagetherebetween, and wherein the suction side outer wall and the suctionside camber line rib define a suction side flow passage therebetween. 5.The blade of claim 1, wherein the camber line rib includes a pressureside camber line rib connected to the pressure side outer wall at apoint aft of the leading edge transverse rib causing the central passageto extend towards the pressure side outer wall.
 6. The blade of claim 5,wherein, at a location where the central passage extends towards thepressure side outer wall, a width t_(f) is defined between outermostextents of the central passage and a width t_(a) is defined betweenouter surfaces of the suction side outer wall and the pressure sideouter wall, and wherein a ratio of the width t_(f) to the width t_(a)ranges from 40% to 70%.
 7. The blade of claim 1, wherein the camber linerib includes a suction side camber line rib connected to the suctionside outer wall at a point aft of the leading edge transverse ribcausing the central passage to extend towards the suction side outerwall.
 8. The blade of claim 7, wherein, at a location where the centralpassage extends towards the suction side outer wall, a width t_(f) isdefined between outermost extents of the central passage and a widtht_(a) is defined between outer surfaces of the suction side outer walland the pressure side outer wall, and wherein a ratio of the width t_(f)to the width t_(a) ranges from 40% to 70%.
 9. The blade of claim 1,wherein the leading edge transverse rib includes a crossover passagebetween the leading edge passage and the central passage.
 10. The bladeof claim 1, wherein the camber line rib has a wavy profile.
 11. Theblade of claim 1, wherein the blade comprises one of a turbine rotorblade or a turbine stator blade.
 12. A turbine rotor blade comprising anairfoil defined by a concave pressure side outer wall and a convexsuction side outer wall that connect along leading and trailing edgesand, therebetween, form a radially extending chamber for receiving theflow of a coolant, the turbine rotor blade further comprising: a ribconfiguration including: a leading edge transverse rib connecting thepressure side outer wall and the suction side outer wall andpartitioning the radially extending chamber into a leading edge passagewithin the leading edge of the airfoil and a central passage adjacent tothe leading edge passage; and a camber line rib connecting to a selectedone of the pressure side outer wall or the suction side outer wall at apoint aft of the leading edge transverse rib causing the central passageto extend towards the selected one of the pressure side outer wall orthe suction side outer wall.
 13. The turbine rotor blade of claim 12,wherein the camber line rib includes: a pressure side camber line ribresiding near the pressure side outer wall and connected to the pressureside outer wall at a point aft of the leading edge transverse ribcausing the central passage to extend towards the pressure side outerwall; and a suction side camber line rib residing near the suction sideouter wall and connected to the suction side outer wall at a point aftof the leading edge transverse rib causing the central passage to extendtowards the suction side outer wall.
 14. The turbine rotor blade ofclaim 13, wherein, at a location where the central passage extendstowards the suction side outer wall and the pressure side outer wall, awidth t_(f) is defined between outermost extents of the central passageand a width t_(a) is defined between outer surfaces of the suction sideouter wall and the pressure side outer wall, and wherein a ratio of thewidth t_(f) to the width t_(a) ranges from 40% to 70%.
 15. The turbinerotor blade of claim 13, wherein the pressure side outer wall and thepressure side camber line rib define a pressure side flow passagetherebetween, and wherein the suction side outer wall and the suctionside camber line rib define a suction side flow passage therebetween.16. The turbine rotor blade of claim 12, wherein the camber line ribincludes a pressure side camber line rib connected to the pressure sideouter wall at a point aft of the leading edge transverse rib causing thecentral passage to extend towards the pressure side outer wall.
 17. Theturbine rotor blade of claim 12, wherein the camber line rib includes asuction side camber line rib connected to the suction side outer wall ata point aft of the leading edge transverse rib causing the centralpassage to extend towards the suction side outer wall.
 18. The turbinerotor blade of claim 12, wherein the leading edge transverse ribincludes a crossover passage between the leading edge passage and thecentral passage.
 19. The turbine rotor blade of claim 12, wherein thecamber line rib has a wavy profile.